Gas turbine combustion acoustic damping system

ABSTRACT

A gas turbine may include a rotatable shaft, a compressor disposed about the rotatable shaft and configured to output compressed air, and a combustor disposed about the rotatable shaft. The combustor may be configured to receive the compressed air and output high temperature compressed gas. The gas turbine may further include a power turbine disposed about the rotatable shaft and configured to receive the high temperature compressed gas, and a first liner defining a plurality of holes and disposed around the combustor. The power turbine may be configured to expand the high temperature compressed gas and rotate the rotatable shaft. The first liner may have a first end and a longitudinally opposite second end. The first end may be coupled to an inner surface of the casing at or adjacent an upstream end of the combustor and the second end may be substantially free from any connection with the casing.

CROSS-REFERENCE TO RELATED APPLICATIONS

The present application is a continuation of co-pending U.S. patentapplication having Ser. No. 14/566,963, filed on Dec. 11, 2014, whichclaims the benefit of U.S. Provisional Patent Application having Ser.No. 61/940,956, filed on Feb. 18, 2014. These priority applications arehereby incorporated by reference in their entirety into the presentapplication to the extent consistent with the present application.

BACKGROUND

Gas turbines may include a compressor for compressing air, a combustorfor producing a hot gas by burning fuel in the presence of thecompressed air produced by the compressor, and a turbine for expandingthe hot gas to extract shaft power. The combustor may be operated suchthat a low level of emissions, such as oxides of nitrogen (NOx), areproduced by the combustor.

In order to reduce the amount of NOx emissions, a lean-premix fuel maybe provided to the combustor. A fuel-lean premix may include fuelpremixed with an excess of air (e.g., air in a quantity more thanstoichiometrically required for combustion). While the fuel-lean premixmay reduce the amount of NOx emissions, high frequency combustioninstabilities, commonly referred to as “high frequency dynamics” or“screech oscillations,” may result from burning rate fluctuations insidethe combustors when the fuel-lean premix is burned in the combustors.These burning rate fluctuation instabilities may create pressure waves(also referred to as acoustic energy) that may damage the combustor.

One way to reduce these damaging pressure waves is to operate thecombustor at maximum power and at standard atmospheric conditions, anddesign the combustor such that the frequency of pressure waves does notcoincide with the natural frequency of oscillation of the sheet metal ofthe combustor. However, gas turbines may generally operate to provide awide range of output power under a wide range of operating temperatureand pressure, and, as a result, pressure waves having a range offrequencies may be generated in the combustor. It may, therefore, bedifficult to design a combustor such that the frequency of pressurewaves does not coincide with the natural frequency of oscillation of thesheet metal of the combustor.

What is needed, then, is a combustor of a gas turbine that may producelow level NOx emissions and dampen the generated acoustic energy whileoperating over a wide range of operating temperatures and pressures.

SUMMARY

Example embodiments of the disclosure may provide a gas turbine. The gasturbine may include a rotatable shaft, a compressor disposed about therotatable shaft and configured to output compressed air, and a combustordisposed about the rotatable shaft and at least partially enclosed in acasing of the gas turbine. The combustor may be configured to receivethe compressed air and output high temperature compressed gas having atemperature greater than the compressed air. The gas turbine may furtherinclude a power turbine disposed about the rotatable shaft andconfigured to receive the high temperature compressed gas from thecombustor, and a first liner that defines a plurality of holes and isdisposed around the combustor. The power turbine may be configured toexpand the high temperature compressed gas and rotate the rotatableshaft. The first liner may have a first end and a longitudinallyopposite second end. The first end may be coupled to an inner surface ofthe casing at or adjacent an upstream end of the combustor and thesecond end may be substantially free from any connection with thecasing.

Example embodiments of the disclosure may provide another gas turbine.The gas turbine may include a rotatable shaft, a compressor disposedabout the rotatable shaft and configured to output compressed air, and acombustor disposed about the rotatable shaft and at least partiallyenclosed in a casing of the gas turbine. The combustor may be configuredto receive the compressed air and to output high temperature compressedgas having a temperature greater than the compressed air. The compressedair and fuel may be ignited in a combustion zone of the combustor toproduce the high temperature compressed gas. The combustor may define aplurality of effusion cooling holes disposed adjacent the combustionzone. The gas turbine may further include a power turbine disposed aboutthe rotatable shaft and configured to receive the high temperaturecompressed gas from the combustor, and a first liner defining aplurality of holes and disposed around the combustor. The power turbinemay be configured to expand the high temperature compressed gas androtate the rotatable shaft. The first liner may have a first end and alongitudinally opposite second end. The first end may be coupled to aninner surface of the casing at or adjacent an upstream end of thecombustor and the second end may be substantially free from anyconnection with the casing.

Example embodiments of the disclosure may provide yet another gasturbine. The gas turbine may include a rotatable shaft, a compressordisposed about the rotatable shaft and configured to output compressedair, a combustor disposed about the rotatable shaft and configured tooutput high temperature compressed gas having a temperature greater thanthe compressed air, and a power turbine disposed about the rotatableshaft and configured to receive the high temperature compressed gas fromthe combustor. The power turbine may be configured to expand the hightemperature compressed gas and rotate the rotatable shaft. The gasturbine may further include a first liner defining a plurality ofimpingement cooling holes and a second acoustic liner defining aplurality of holes and having a first end and a longitudinally oppositesecond end. The first end of the second acoustic liner may be coupled tothe first liner. The first liner may be coupled to an outer surface ofthe combustor and the first liner and the outer surface of the combustormay define an acoustic chamber therebetween. At least a portion of theouter surface defining the acoustic chamber may define a plurality ofeffusion cooling holes, and the plurality of impingement cooling holesand the plurality of effusion cooling holes may be configured toattenuate acoustic energy generated in the combustor.

BRIEF DESCRIPTION OF THE DRAWINGS

The present disclosure is best understood from the following detaileddescription when read with the accompanying Figures. It is emphasizedthat, in accordance with the standard practice in the industry, variousfeatures are not drawn to scale. In fact, the dimensions of the variousfeatures may be arbitrarily increased or reduced for clarity ofdiscussion.

FIG. 1 is a schematic diagram of a gas turbine including a combustor,according to example embodiments.

FIG. 2 is a partial cross-sectional view illustrating in relativelygreater detail the gas turbine including the combustor of FIG. 1.

FIG. 3A illustrates a plan view of the combustor in FIGS. 1 and 2, thecombustor including a swirler, a mixer, and a combustor body, accordingto example embodiments.

FIG. 3B illustrates a cross-sectional view of the combustor in FIG. 3Aalong line 3B-3B, according to example embodiments.

FIG. 3C illustrates a cross-sectional view of the swirler in FIG. 3Aalong line 3C-3C, according to example embodiments.

FIG. 4 illustrates a cross-sectional view of the mixer and combustorbody in FIGS. 3A and 3B disposed in an acoustic liner in the gas turbineof FIG. 1, according to example embodiments.

FIG. 5 is a partial cross-sectional view illustrating in relativelygreater detail the gas turbine in FIG. 1, the gas turbine including thecombustor of FIGS. 1, 2, 3A, and 3B disposed in a first acoustic linerand a second liner, according to example embodiments.

DETAILED DESCRIPTION

It is to be understood that the following disclosure describes severalexemplary embodiments for implementing different features, structures,or functions of the present disclosure. Exemplary embodiments ofcomponents, arrangements, and configurations are described below tosimplify the present disclosure; however, these exemplary embodimentsare provided merely as examples and are not intended to limit the scopeof the present disclosure. Additionally, the present disclosure mayrepeat reference numerals and/or letters in the various exemplaryembodiments and across the Figures provided herein. This repetition isfor the purpose of simplicity and clarity and does not in itself dictatea relationship between the various exemplary embodiments and/orconfigurations discussed in the various Figures. Moreover, the formationof a first feature over or on a second feature in the description thatfollows may include embodiments in which the first and second featuresare formed in direct contact, and may also include embodiments in whichadditional features may be formed interposing the first and secondfeatures, such that the first and second features may not be in directcontact. Finally, the exemplary embodiments presented below may becombined in any combination of ways, i.e., any element from oneexemplary embodiment may be used in any other exemplary embodiment,without departing from the scope of the disclosure.

Additionally, certain terms are used throughout the followingdescription and the claims to refer to particular components. As oneskilled in the art will appreciate, various entities may refer to thesame component by different names, and as such, the naming conventionfor the elements described herein is not intended to limit the scope ofthe present disclosure, unless otherwise specifically defined herein.Further, the naming convention used herein is not intended todistinguish between components that differ in name but not function.Additionally, in the following discussion and in the claims, the terms“including” and “comprising” are used in an open-ended fashion, and thusshould be interpreted to mean “including, but not limited to.” Allnumerical values in this disclosure may be exact or approximate valuesunless otherwise specifically stated. Accordingly, various embodimentsof the disclosure may deviate from the numbers, values, and rangesdisclosed herein without departing from the intended scope. Furthermore,as it is used in the claims or specification, the term “or” is intendedto encompass both exclusive and inclusive cases, i.e., “A or B” isintended to be synonymous with “at least one of A and B,” unlessotherwise expressly specified herein.

FIG. 1 is a schematic diagram of a gas turbine 100, according to exampleembodiments. The gas turbine 100 may include a compressor 102 (e.g., acentrifugal compressor) that may be driven by a power turbine 106 via arotatable shaft 101. Ambient air 112 may be drawn into the compressor102 and may be compressed. The compressed air 108 may be directed to acombustion system that may include a plurality of combustors 104 (e.g.,can-annular combustors, only one combustor 104 is illustrated)circumferentially disposed around the shaft 101 and a fuel nozzle 118that may introduce fuel 116 from an external fuel source (not shown)into the plurality of combustors 104. In the combustors 104, the fuel116 may be burned in the compressed air 108 to produce a hightemperature compressed gas 110 having a temperature relatively greaterthan the compressed air 108.

The high temperature compressed gas 110 produced by the combustors 104may be directed to the power turbine 106 where high temperaturecompressed gas 110 may be expanded, thereby producing shaft power fordriving the compressor 102. The expanded gas 114 produced by the powerturbine 106 may be exhausted, for example, to the atmosphere.

FIG. 2 is a partial cross-sectional view illustrating in relativelygreater detail the gas turbine 100 including the combustor 104. Althoughonly one combustor 104 is shown, it will be understood that the gasturbine 100 may include a plurality of combustors 104 circumferentiallydisposed about the shaft, generally indicated at 101. The plurality ofcombustors 104 may be enclosed by an external casing 122 of the gasturbine 100. The compressed air 108 may enter an area between the casing122 and the combustor 104, as generally indicated by the arrow A. Eachcombustor 104 may define a combustion zone 130 (e.g., a single combustorzone) located therein and a mixture of the compressed air 108 and fuel116 may be ignited in the combustion zone 130. Due to combustion of thecompressed air 108 and fuel 116, the high temperature compressed gas 110may be produced and may be directed to the power turbine 106 (FIG. 1),as generally indicated by the arrow B.

FIG. 3A illustrates a plan view of the combustor 104 including a swirler302, according to example embodiments. FIG. 3B illustrates across-sectional view of the combustor 104 along line 3B-3B in FIG. 3A,according to example embodiments. FIG. 3C illustrates a cross-sectionalview of the swirler 302 of the combustor 104 along line 3C-3C in FIG.3A, according to example embodiments. Referring to FIGS. 3A and 3B, thecombustor 104 may include a swirler 302 at an upstream end 318 and agenerally cylindrical combustor body 306 at a downstream end 320. Theswirler 302 may be in fluid communication with the combustor body 306via a generally cylindrical mixer 304 disposed therebetween. Asillustrated, a radial extent of the mixer 304 may be relatively smallerthan a radial extent of the swirler 302 and a radial extent of thecombustor body 306. The swirler 302, the mixer 304, and the combustorbody 306 may be coupled to each other in an end-to-end relationshipalong a central axis 301 of the combustor 104. For example, an inlet(not shown) of the mixer 304 may be coupled to an outlet (not shown) ofthe swirler 302 and an outlet of the mixer 304 may be coupled to aninlet (not shown) of the combustor body 306.

A first acoustic liner 308 may be disposed encircling a portion of anouter surface 307 of the combustor body 306. The first acoustic liner308 may be radially spaced from the outer surface 307 and may be coupledthereto via sidewalls 308 a, 308 b extending from the edges of the firstacoustic liner 308. As shown, the first acoustic liner 308 may generallybe disposed at or adjacent the location of the combustion zone 130 inthe combustor body 306. An acoustic chamber 310 may be formed by thefirst acoustic liner 308 and the outer surface 307 connected to eachother via the sidewalls 308 a, 308 b. A plurality of circumferentiallydisposed dilution holes 311 may be defined by the combustor body 306 ator adjacent the downstream end 320. The high temperature compressed gas110 (FIG. 1) may exit from the downstream end 320.

Referring to FIG. 3C in conjunction with FIGS. 1 and 3B, the swirler 302may include a plurality of swirler vanes 312 disposed about the centralaxis 301 (FIG. 3B). A plurality of swirler channels 313 may be definedby the swirler 302, each swirler channel 313 being disposed betweenadjacent swirler vanes 312. Fuel 116 (FIG. 1) may enter the combustor104 via fuel pegs 314 disposed in a generally circular manner about theswirler vanes 312. As will be described later, pilot fuel may beintroduced in the swirler 302 via pilot fuel holes 316. The compressedair 108 (FIG. 1) may enter the swirler 302 and may traverse the swirlerchannels 313, and the swirler vanes 312 may impart a swirling motion tothe compressed air 108 traversing the swirler channels 313. The swirlingcompressed air 108 and fuel 116 may flow downstream into the mixer 304and mix with each other in the mixer 304. The swirling compressed air108 and fuel 116 may continue to mix with each other as the swirlingcompressed air 108 and fuel 116 travel further downstream into thecombustor body 306. The mixture of fuel 116 and swirling compressed air108 may be ignited in the combustion zone 130. The swirling motionimparted by the swirler vanes 312 to the compressed air 108 may ensurerelatively uniform mixing of the fuel 116 and the compressed air 108,thereby reducing locally fuel-rich mixtures and the associated hightemperatures that may increase NOx generation.

FIG. 4 illustrates a cross-sectional view of the mixer 304 and combustorbody 306 of the combustor 104 disposed in the first acoustic liner 308in the gas turbine 100, according to example embodiments. A plurality ofeffusion cooling holes 322 may be formed in the portion of the outersurface 307 that at least partially defines the acoustic chamber 310.The plurality of effusion cooling holes 322 may be through holes and maybe disposed in the form of a matrix on the outer surface 307. The firstacoustic liner 308 may define a plurality of impingement cooling holes324. The plurality of impingement cooling holes 324 may be disposed inthe form of a matrix on the first acoustic liner 308. In an exampleembodiment, the impingement cooling holes 324 may be relatively smallerin size, e.g., diameter, than the effusion cooling holes 322. Theeffusion cooling holes 322 and the impingement cooling holes 324 may begenerally circular in cross section; however, other shapes, such ashexagonal or elliptical, are also envisioned without departing from thescope of the present disclosure.

A generally cylindrical and longitudinally extending second liner 326may be disposed about the combustor body 306 and the first acousticliner 308. The second liner 326 may or may not be an acoustic liner. Thesecond liner 326 may include two or more segments (326 a, 326 b)connected to each other in an end-to-end relationship. As illustrated, afirst segment 326 a of the second liner 326 may define the plurality ofopenings or holes 328 disposed along a length thereof. Although notillustrated, a second segment 326 b may also define a plurality ofholes. A first end 327 of the second liner 326 (also referred to as theupstream end of the first segment 326 a) may be coupled to an innersurface 124 of the casing 122 at or adjacent the upstream end 318 of thecombustor 104, while a longitudinally opposite second end 329 (alsoreferred to as the downstream end of the second segment 326 b) may bedisposed at or adjacent the downstream end 320 of the combustor 104. Forexample, referring to FIGS. 2 and 4, and with respect to the axis ofrotation (not shown) of the shaft 101, a radially outer portion 329 a ofthe second liner 326 at the second end 329 may be substantially freefrom any connection with the casing 122, while a radially inner portion329 b of the second liner 326 at the second end 329 may be coupled at oradjacent the downstream end 320 of the combustor 104. In an exampleembodiment, the radially outer portion 329 a of the second liner 326 mayextend generally circumferentially around some or all the plurality ofcombustors 104 of the gas turbine 100, thereby enclosing (at leastpartially) the plurality of combustors 104. As used in the presentdisclosure, the term “substantially free from any connection” may referto a component completely lacking a connection or having a minimalconnection with another component.

Referring to FIG. 4 in conjunction with FIGS. 1, 2, 3A, 3B, and 3C, thecompressed air 108 may be received from the compressor 102, as generallyindicated by the arrow A. The compressed air 108 may enter an area 303between the combustor 104 and the second liner 326 via the plurality ofholes 328 in the second liner 326, as generally indicated by arrows C.The compressed air 108 flowing across the second liner 326 may besubjected to a relatively low pressure drop (e.g., about 3% to about4%). The compressed air 108 may impinge on the outer surface 307 of thecombustor 104 and the first acoustic liner 308 (thereby cooling theouter surface 307 and the first acoustic liner 308). A portion (e.g.,about 30% to about 40%) of the compressed air 108, generally indicatedby the arrows F, may enter the swirler 302 and traverse the swirlerchannels 313, mix with the fuel 116 in the mixer 304 (resulting in thepre-mixed/pre-swirled mixture, discussed above), and enter the combustorbody 306 where the mixture of the fuel 116 and compressed air 108 may beignited to produce the high temperature compressed gas 110. Anotherportion (e.g., about 30% to about 40%) of the compressed air 108 mayenter the combustor body 306 via the dilution holes 311 and may mix withthe high temperature compressed gas 110 exiting the combustor body 306.The compressed air 108 entering via the dilution holes 311 may berelatively cooler than the high temperature compressed gas 110 and maylower the temperature of the high temperature compressed gas 110 to alevel suitable for supplying to the power turbine 106.

Some of the remaining compressed air 108 in the area 303 may enter theacoustic chamber 310 through the impingement cooling holes 324, asgenerally indicated by arrows D, while the rest may enter through holes(not shown) defined in the second end 329 of the second line 326. Thecompressed air 108 entering the acoustic chamber 310 may be accelerateddue to the relatively smaller size of the impingement cooling holes 324.The compressed air 108 may also experience a relatively greater pressuredrop as the compressed air 108 flows across the first acoustic liner308. As a result, the acoustic chamber 310 may include air having asubstantially reduced pressure (about 1% to about 2%) relative to thecompressed air 108 from the compressor 102.

Due to the relatively larger size of the effusion cooling holes 322, thelow pressure air in the acoustic chamber 310 may enter the combustorbody 306 (at or adjacent the combustion zone 130) with a substantiallyreduced velocity compared to the velocity of the compressed air 108 whenentering the acoustic chamber 310. As shown in FIG. 4, the low pressureair entering the combustor body 306, generally indicated by the arrowsE, may flow along the inner surface of the combustor body 306. The lowpressure air entering the combustion zone 130 may be minimized and thismay avoid localized cooling (also referred to as quenching) and preventformation of gases, such as, carbon monoxide (CO).

In operation, a flame (not shown) may be initially established in themixer 304 and/or the combustion zone 130 by the introduction of a smallquantity of fuel 116 (also referred to as pilot fuel), via the pilotfuel holes 316 in the swirler 302. As an increase in the load on thepower turbine 106 may require an increased output from the combustor104, additional fuel 116 may be added via the fuel pegs 314. Since thefuel pegs 314 may result in a better distribution of the fuel 116 withinthe compressed air 108, a leaner fuel/air mixture may be produced, whichmay reduce NOx generation. Once ignition is established in thecombustion zone 130 and the gas turbine 100 has reached a predeterminedpower output level, e.g., greater than about 80% of full power, the fuel116 to the pilot fuel holes 316 may be shut-off and the gas turbine 100may operate under premix (leaner fuel/air mixture) combustion.

FIG. 5 is a partial cross-sectional view illustrating in relativelygreater detail the gas turbine 100, the gas turbine 100 including thecombustor 104 disposed in a first acoustic liner 308 and a second liner326, according to example embodiments. The gas turbine 100, and thecombustor 104 disposed in a first acoustic liner 308 and a second liner326 as illustrated in FIG. 5 may be similar in some respects to the gasturbine, and the combustor disposed in the first acoustic liner and thesecond liner in FIGS. 2 and 4 described above and therefore may be bestunderstood with reference to the description of FIGS. 2 and 4 where likenumerals designate like components and will not be described again indetail. As illustrated in FIG. 5, the first acoustic liner 308 may bedisposed encircling a portion of the outer surface 307 of the combustorbody 306 and the first acoustic liner 308 may be generally disposed ator adjacent the location of the combustion zone 130 in the combustorbody 306. The first acoustic liner 308 may be radially spaced from theouter surface 307 and an upstream end (not shown) of the first acousticliner 308 may be coupled to the mixer 304, generally at 305. Adownstream end (not shown) of the first acoustic liner 308 may becoupled to the outer surface 307 of the combustor body 306 via asidewall 309. The first acoustic liner 308 may generally follow theouter surface 307 (e.g., disposed generally parallel to the outersurface 307) of the combustor body 306. Although not illustrated, thefirst acoustic liner 308 may also be radially spaced from an outersurface (or at least a portion thereof) of the mixer 304 and may becoupled to the outer surface of the mixer 304 at or adjacent an upstreamend thereof.

As illustrated, an acoustic chamber 310 may be defined by the firstacoustic liner 308 and the outer surface 307 connected to each other.The first acoustic liner 308 may define the plurality of impingementcooling holes 324. The portion of the combustor 104 adjacent the firstacoustic liner 308 may define the plurality of effusion cooling holes322 adjacent the combustion zone 130. The size (e.g., diameter) of theeffusion cooling holes 322 may be greater than the size (e.g., diameter)of the impingement cooling holes 324.

The second liner 326 may be generally cylindrical in shape and mayextend longitudinally along the combustor body 306. The second liner 326may comprise two or more segments (326 a, 326 b) connected to each otherin an end-to-end relationship. As illustrated, the first segment 326 aof the second liner 326 may define the plurality of openings or holes328 disposed along a length thereof. Although not illustrated, thesecond segment 326 b may also define a plurality of holes. The first end327 of the second liner 326 (also referred to as the upstream end of thefirst segment 326 a) may be coupled to the first acoustic liner 308,while the longitudinally opposite second end 329 (also referred to asthe downstream end of the second segment 326 b) may be disposed at oradjacent the downstream end 320 of the combustor 104. As illustrated,with reference to the axis of rotation (not shown) of the shaft 101, aradially outer portion 329 a of the second liner 326 at the second end329 may not be coupled to the combustor 104, while a radially innerportion 329 b of the second liner 326 at the second end 329 may becoupled at or adjacent the downstream end 320 of the combustor 104. Inan example embodiment, the radially outer portion 329 a of the secondliner 326 may extend generally circumferentially around some or all ofthe plurality of combustors 104 of the gas turbine 100, therebyenclosing (at least partially) the plurality of combustors 104. As such,only the first end 327 of the second liner 326 may be coupled to thefirst acoustic liner 308.

The compressed air 108 (FIG. 1) from the compressor 102 (FIG. 1) may bereceived in an area between the casing 122 and the combustor 104, asgenerally indicated by the arrow A. A first portion (e.g., about 30% toabout 40%) of the compressed air 108 may flow across the second liner326 via the plurality of holes 328, as generally indicated by arrows C.The first portion of the compressed air 108 may enter the combustor body306 via the dilution holes 311 (not shown) and may mix with the hightemperature compressed gas 110 (generated by the combustion of the fuel116 and the compressed air 108) exiting the combustor body 306. Thefirst portion of the compressed air 108 entering via the dilution holes311 may be relatively cooler than the high temperature compressed gas110 and may lower the temperature of the high temperature compressed gas110 to a level suitable for supplying to the power turbine 106 (FIG. 1).A second portion (e.g., about 30% to about 40%) of the compressed air108, indicated by the arrows F, may enter the swirler 302 (FIGS. 3A-3C)and traverse the swirler channels 313 in the swirler 302, mix with thefuel 116 in the mixer 304 (FIGS. 3A-3C), and enter the combustor body306 where the mixture of the fuel 116 and compressed air 108 may beignited to produce the high temperature compressed gas 110.

Some of the remaining compressed air 108 may enter the acoustic chamber310 through the impingement cooling holes 324 on the first acousticliner 308, as generally indicated by arrows D, while the rest of thecompressed air 108 may enter through holes (not shown) defined in thesecond end 329 of the second liner 326. The compressed air 108 enteringthe acoustic chamber 310 may be accelerated due to the relativelysmaller size of the impingement cooling holes 324 and may experience arelatively greater pressure drop (e.g., greater differential pressureacross the first acoustic liner 308) as the compressed air 108 flowsacross the first acoustic liner 308. As a result, the acoustic chamber310 may include air having a substantially reduced pressure relative tothe compressed air 108 received from the compressor 102.

The low pressure air in the acoustic chamber 310 may enter the combustorbody 306 at or adjacent the combustion zone 130 via the plurality ofeffusion cooling holes 322. The low pressure air entering the combustorbody 306 may be generally indicated by the arrows E. The low pressureair may be prevented from entering the combustion zone 130, therebyavoiding localized cooling (also referred to as quenching) andpreventing formation of gases, such as, carbon monoxide (CO). Asindicated by the arrows E in FIG. 5, the low pressure air may flow alongthe inner surface of the combustor 104 and may form a film on the innersurface.

The operation of the gas turbine 100 including the combustor 104illustrated in FIG. 5 may be similar to the operation of the gas turbine100 including the combustor 104 described with respect to FIGS. 2, 3A-3Cand 4, and will, therefore, be omitted for the sake of brevity.

According to example embodiments illustrated in FIGS. 2, 3A-3C, 4, and5, the first acoustic liner 308 with the impingement cooling holes 324and the combustor body 306 with the effusion cooling holes 322 mayfunction as a Helmholtz oscillator that may dampen the acoustic energy,e.g., due to the damaging pressure waves, in the combustor 104. Examplesof acoustic liners functioning as a Helmholtz oscillator may be found inco-owned U.S. Pat. No. 6,601,672 entitled “Double Layer Acoustic Linerand a Fluid Pressurizing Device and Method Utilizing Same” and co-ownedU.S. Pat. No. 8,596,413 entitled “Acoustic Array of Polymer Material,”each of these co-owned patents are incorporated herein by reference tothe extent not inconsistent with the present disclosure. According toexample embodiments, the temperature, pressure, and/or relativegeometries of the impingement cooling holes 324, the effusion coolingholes 322, and/or the acoustic chambers 310 may determine the acousticfrequencies generated in the combustor 104 that may be damped.

The foregoing has outlined features of several embodiments so that thoseskilled in the art may better understand the present disclosure. Thoseskilled in the art should appreciate that they may readily use thepresent disclosure as a basis for designing or modifying other processesand structures for carrying out the same purposes and/or achieving thesame advantages of the embodiments introduced herein. Those skilled inthe art should also realize that such equivalent constructions do notdepart from the spirit and scope of the present disclosure, and thatthey may make various changes, substitutions, and alterations hereinwithout departing from the spirit and scope of the present disclosure.

We claim:
 1. A gas turbine, comprising: a rotatable shaft; a compressordisposed about the rotatable shaft and configured to output compressedair; a can-annular combustion system comprising a plurality ofcombustors disposed about the rotatable shaft and at least partiallyenclosed in a casing of the gas turbine, a combustor of the plurality ofcombustors configured to receive the compressed air and output hightemperature compressed gas having a temperature greater than atemperature of the compressed air; a power turbine disposed about therotatable shaft and configured to receive the high temperaturecompressed gas from the combustor, the power turbine configured toexpand the high temperature compressed gas and rotate the rotatableshaft; and an acoustic liner coupled to an outer surface of thecombustor and defining an acoustic chamber between the acoustic linerand the outer surface of the combustor, the acoustic liner defining aplurality of impingement cooling holes and being configured to attenuateacoustic energy generated in the combustor, wherein a portion of theouter surface of the combustor that defines the acoustic chamber furtherdefining a plurality of effusion cooling holes, wherein a diameter ofeach effusion cooling hole of the plurality of effusion cooling holes isgreater than a diameter of each impingement cooling hole of theplurality of impingement cooling holes; a second liner extending along alongitudinal axis of the combustor and around the combustor, the secondliner defining a plurality of holes and having a first end and alongitudinally opposite second end, the first end being coupled to aninner surface of the casing of the gas turbine at or adjacent anupstream end of the combustor and the longitudinally opposite second endbeing free from a connection with the casing of the gas turbine, whereinthe compressed air from the compressor entering the acoustic chamberthrough the plurality of impingement cooling holes, due to therelatively smaller diameter of the plurality of impingement coolingholes, experiences a pressure drop as the compressed air flows acrossthe plurality of impingement cooling holes in the acoustic liner,thereby the acoustic chamber being with air having a reduced pressurecompared to a pressure of the compressed air from the compressor priorto flowing through the plurality of impingement cooling holes, whereinthe reduction of pressure of air in the acoustic chamber is in a rangefrom 1% to 2% relative to the pressure of the compressed air from thecompressor prior to flowing through the plurality of impingement coolingholes, wherein air exits the acoustic chamber through the plurality ofeffusion cooling holes to form a cooling film of air along an innersurface of the combustor, wherein, based on the reduced pressure of airin the acoustic chamber, an amount of the exiting air that enters into acombustion zone of the combustor is reduced, and this reduction iseffective to avoid localized cooling and formation of undesirable gasesin the combustion zone.
 2. The gas turbine of claim 1, wherein adifferential pressure across the acoustic liner is substantially greaterthan a differential pressure across a portion of the combustor adjacentthe acoustic liner.
 3. The gas turbine of claim 1, wherein adifferential pressure across the acoustic liner is substantially greaterthan a differential pressure across the second liner.
 4. The gas turbineof claim 1, wherein the second liner is disposed such that thecompressed air is received between the casing of the gas turbine and thesecond liner, the compressed air flowing through the plurality of holesin the second liner.
 5. The gas turbine of claim 4, wherein the acousticliner is disposed such that at least a portion of the compressed airflowing across the second liner is input to the acoustic chamber via theplurality of impingement cooling holes.
 6. The gas turbine of claim 4,wherein the combustor defines a plurality of dilution holescircumferentially disposed at or adjacent a downstream end thereof andat least a portion of the compressed air flowing across the second lineris input to the combustor via the plurality of dilution holes.
 7. Thegas turbine of claim 4, wherein at least a portion of the compressed airflowing across the second liner is input to the combustor at theupstream end of the combustor.
 8. The gas turbine of claim 7, whereinthe combustor comprises a swirler disposed at the upstream end thereof,the swirler configured to receive the compressed air, the swirlerincluding a plurality of swirler vanes configured to swirl thecompressed air.
 9. The gas turbine of claim 8, wherein: the combustorfurther comprises a mixer and a combustor body, the swirler being influid communication with the combustor body via the mixer, the mixerbeing disposed between the swirler and the combustor body, the mixer isconfigured to mix fuel with the swirled compressed air received from theswirler prior to outputting a mixture of the fuel and the swirledcompressed air to the combustor body, and the mixture of the fuel andthe swirled compressed air is ignited in a combustion zone defined inthe combustor body.
 10. The gas turbine of claim 9, wherein theplurality of effusion cooling holes is disposed adjacent the combustionzone.